The present invention relates generally to maintaining gas turbine engines, and more particularly, to maintaining turbine discs prone to cracks in bucket attachment dovetails.
As shown in FIG. 1, a gas turbine 10 may have a combustion section 12 in a gas flow path between a compressor 14 and a turbine 16. The combustion section 12 may include an annular array of combustion chambers 20, e.g., combustion cans. The turbine 16 may be coupled to rotationally drive the compressor 14 and a power output drive shaft 18. Air may enter the gas turbine 10 and pass through the compressor 14. Temperatures in the compressor 14 may range from ambient temperature to about 800° F. Pressures in the compressor 14 may range from local atmospheric pressure to about 228 pounds per square inch (“psi”). High pressure air from the compressor 14 may enter the combustion section 12 where it may mix with fuel and burn. High energy combustion gases may exit the combustion section 12 to power the turbine 16 which, in turn, may drive the compressor 14 and the output power shaft 18. Temperatures in the turbine 16 may range from about 2540° F. at the inlet to about 1200° F. at the exhaust. Pressures in the turbine 16 may range from about 15 psi to about 200 psi. The extreme temperature and pressure environment of a turbine may present a number of challenges for manufacture and maintenance of turbine components.
The turbine 16 of the gas turbine 10 typically has multiple sets or stages of stationary blades, known as nozzles or vanes, and moving blades, known as rotor blades or buckets. At each stage, the buckets (not shown) may be mounted on a turbine disc 22, as illustrated in FIG. 2. The buckets may attach to the turbine disc 22 with dovetails 24. As illustrated in FIG. 3, the stages of turbine discs 22 may be aligned coaxially. The first stage of the turbine 16 is the section immediately adjacent to the combustion section 12 of the gas turbine 10 and, thus, is the region of the turbine 16 that is exposed to the highest temperatures. Similarly, the second stage of the turbine 16 is the section immediately adjacent to the first stage, and so forth for each stage.
High thermal gradients may exist from the dovetails 24 on the radial perimeter of the turbine disc 22 to the interior regions of the turbine disc 22, with the most severe gradients present in the first stage. Unaddressed, such thermal gradients may create metal fatigue, deformations, and/or cracks in the turbine disc 22. Therefore, as seen in FIG. 4, cooling slots 26 may extend from the interior regions of the turbine disc 22 through the dovetails 24 to provide thermal relief at the dovetail regions of the turbine disc 22.
FIG. 4A illustrates the dovetail region for a typical first stage turbine disc, while FIG. 4B illustrates the same for a typical second stage turbine disc. As shown in FIG. 5, a known problem with turbine discs 22 comprising such cooling slots 26 is cracking 28 at the cooling slots 26. It is believed in the industry that such cracking is not amenable to repair. With tensile stresses and temperatures during normal operation, crack initiation can expose the alloy of the disc 22 to oxidation along its grain boundaries. This phenomenon is known as stress-accelerated grain boundary oxidation (“SAGBO”). With continued exposure, this crack propagation may result in liberation of the bucket attachment dovetail and the bucket, causing extensive hot gas path damage. The estimated costs of catastrophic disc failure can exceed $10 million in direct equipment damage alone, with the gross damage estimates as high as $20 million. Therefore, identification of cracks in the dovetail region has heretofore necessitated replacement of the turbine disc 22. As has been experienced in the industry, such cracking 28 in the dovetail region may reduce the useful life of a disc from approximately 20 years to less than 4 years. In addition to the $6-8 million cost of replacing the turbine disc 22, replacement costs may also include those for plant shutdown and outage, locating and acquiring a replacement disc, and de-stacking and re-stacking the rotor.
Moreover, identification of cracks in the dovetail region may pose challenges. Inspection of this region of the turbine disc 22 has not been typically incorporated into standard maintenance routines, so standard tools and procedures are lacking. Costs of de-stacking and re-stacking the rotor may create a preference for in situ inspection. Visual inspecting for cracking 28 with linear dimensions of less than several thousandths of an inch may require expensive magnification and logging equipment. Additionally, the geometry of the dovetail region may necessitate specialized equipment, such as conforming probes, to perform the inspections.
Actions which have been proposed to help avoid crack initiation may include edge blending, shot-peening, bucket root modifications, and contouring of the slot corners to reduce stress. See “Technical Information Letter 1539-2”, GE Engineering Product Service, Apr. 4, 2006; “Technical Information Letter 1540-2”, GE Engineering Product Service, Apr. 4, 2006; U.S. Pat. No. 5,141,401; and U.S. Patent Application Publication No. 2007/0269316. Shot-peening (also referred to as “peening”) is a process wherein the surface of a workpiece may be impacted by particles or shot. Peening may generate a residual compressive stress in the metal surface, which is thought to improve fatigue resistance. Thus, localized areas of tensile stress, phase transformations, machine and grinding marks, pits, scratches, and the like, may be effectively eliminated from acting as stress concentration points.
In an effort to identify potential cracking damage, an Original Equipment Manufacturer (“OEM”) has suggested that eddy current inspections should be performed at each hot gas path inspection interval and at major inspection intervals. See “Technical Information Letter 1539-2”, GE Engineering Product Service, Apr. 4, 2006; and “Technical Information Letter 1540-2”, GE Engineering Product Service, Apr. 4, 2006. Turbine operators have expressed concerns that the bucket attachment dovetail cracking problem has not been significantly mitigated by the measures proposed to date. This may be due to the fact that the causes and mechanisms of cracking in the bucket attachment dovetail region are not well understood in the industry.
Coatings and fillers have been utilized on various gas turbine engine components to provide thermal barriers, wear resistance, or corrosion protection. Many coatings, which may be suitable for the lower temperature and lower pressure environment of the compressor, may fail in the higher temperature and higher pressure environment of the turbine. Additionally, many coatings may not be suitable for the extreme thermal gradients present in a cooling slot of a turbine bucket attachment dovetail region. Selection and utilization of an appropriate coating or filler would require a thorough understanding of the causes and mechanisms of the cracking in the bucket attachment dovetail region.